With reference to FIG. 1, a ducted fan gas turbine engine generally indicated at 10 has a principal and rotational axis X-X. The engine comprises, in axial flow series, an air intake 11, a propulsive fan 12, an intermediate pressure compressor 13, a high-pressure compressor 14, combustion equipment 15, a high-pressure turbine 16, and intermediate pressure turbine 17, a low-pressure turbine 18 and a core engine exhaust nozzle 19. A nacelle 21 generally surrounds the engine 10 and defines the intake 11, a bypass duct 22 and a bypass exhaust nozzle 23.
The gas turbine engine 10 works in a conventional manner so that air entering the intake 11 is accelerated by the fan 12 to produce two air flows: a first air flow A into the intermediate pressure compressor 14 and a second air flow B which passes through the bypass duct 22 to provide propulsive thrust. The intermediate pressure compressor 13 compresses the air flow A directed into it before delivering that air to the high pressure compressor 14 where further compression takes place.
The compressed air exhausted from the high-pressure compressor 14 is directed into the combustion equipment 15 where it is mixed with fuel and the mixture combusted. The resultant hot combustion products then expand through, and thereby drive the high, intermediate and low-pressure turbines 16, 17, 18 before being exhausted through the nozzle 19 to provide additional propulsive thrust. The high, intermediate and low-pressure turbines respectively drive the high and intermediate pressure compressors 14, 13 and the fan 12 by suitable interconnecting shafts.
As shown in FIG. 2, the high pressure turbine comprises a rotor disc 36 to which a plurality of rotor blades 38 is attached via a disc rim. The rotor blades are in the path of the hot combustion products and thus require cooling to prevent failure. Air flow from the high pressure compressor, flowing through inner cavities of the rotor blades, is used to cool the rotor blades. The air flow 26 and 34 is bled from a final stage 24 of the high pressure compressor to ensure the pressure of the air in the rotor blades is greater than the surrounding combustion products.
The air flow 26 and 34 from the high pressure compressor also contacts a front face 48 of the rotor disc 36. This air flow has a high temperature and high specific entropy, but nonetheless provides sufficient cooling to the turbine blades. However, when contacting the front face of the rotor disc it can heat the rotor disc to undesirable temperatures, resulting in decreased life of the rotor disc, and can cause areas of high peak temperatures on the front face of the rotor disc, leading to large thermal gradients in the rotor disc.
To protect the rotor disc from failure, it is desirable to further cool the rotor disc 36. This can be done by routing cooling air flow to a rear face and a central bore of the rotor disc. However, the high specific entropy of the air from the final stage of the high pressure compressor makes it inefficient for cooling the rear face and the central bore of the turbine disc. Thus air flows 30 and 32 from lower pressure stages of the compressor can be used to cool the rear face and the central bore of the rotor disc. As a result, the rotor disc is in contact with air flows from different stages of the compressor, all of which are at different temperatures. This can create an undesirable thermal environment for the rotor disc, which can experience large thermal gradients and high thermal stresses. Thus the rotor disc can have a reduced life and compromised design. The different air flows also provide little control over the heat transfer coefficients between the rotor disc and surrounding air, leading to poor rotor disc thermal response, and correspondingly poor rotor blade shroud tip clearance.
GB 2420155 suggests an alternative method for supplying cooling air to the turbine blades. In this cooling system, air is bled from a stage of the compressor upstream of the final stage. The air is then passed through a space between the downstream face of the rotor disc and a vaned cover for re-pressurisation. The pressurised air then passes to the turbine blades via the disc rim.
This system can improve the efficiency of the rotor blade cooling, however it still suffers from the problem that air flow from the final stage of the high pressure compressor contacts the upstream face of the rotor disc, while cooler air from the upstream stage of the compressor surrounds other parts of the rotor disc creating undesirably large thermal gradients. In particular, the system does not address the problem of high peak temperatures on the front face of the rotor disc or high thermal stresses in the rotor disc.